1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine blade with tip region cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, a compressor delivers compressed air to a combustor in which the compressed air is burned with a fuel to produce a high temperature gas flow which is then passed through a turbine to convert the energy in the hot gas flow into mechanical work used to drive the rotor shaft of the engine. The engine efficiency can be increased by passing a higher temperature gas into the turbine. However, material properties and cooling limitations limit the turbine inlet temperature.
One method designers can allow for higher turbine inlet temperatures is to provide higher cooling capability with the pressurized cooling air supplied to the turbine airfoils (rotor blades and stator vanes). Since the first stage turbine airfoils are exposed to the highest temperature gas flow, these airfoils require the most cooling,
FIG. 1 shows a prior art first stage turbine blade external heat transfer coefficient profile. As indicated by the figure, the airfoil leading edge, the suction side immediately downstream of the leading edge, as well as the pressure side trailing edge region of the airfoil experience higher hot gas side external heat transfer coefficient than the mid-chord section of the pressure side and downstream of the suction surfaces. In general, the heat load for the airfoil aft section is higher than the forward section.
Another method used to allow for higher turbine inlet temperatures is to apply a TBC (thermal barrier coating) to the airfoils to provide a thermal protection. As the TBC technology improves, more industrial gas turbine (IGT) blades are applied with a thick or low conductivity TBC. With a thicker TBC, the cooling flow demand is reduced. As a result of this, there is not sufficient cooling flow for the design to split the total cooling flow into three flow circuits and utilize the forward flowing serpentine cooling concept of the prior art. Cooling flow for the blade leading and trailing edges has to be combined with the mid-chord flow circuit to form a single 5-pass serpentine flow circuit. However, for a forward flow 5-pass serpentine circuit with total blade cooling flow, BFM (back flow margin, when the airfoil outside pressure is lower higher than the airfoil inside pressure and the hot gas flows into the airfoil cooling circuitry) may become a design issue. In addition, a single 5-pass aft flowing serpentine circuit for a large chord blade design may yield too high of a cooling air temperature when the cooling air reaches to the airfoil trailing edge section. Thus, a loss of cooling potential for the cooling air to achieve a design metal temperature occurs.